[0001] This invention relates to a splitter shroud disposed between radially inner and outer
circumferential rows of airfoils of a rotor blade assembly for an aircraft gas turbine
engine and, more particularly, to the seals and associated seal assembly at annular
leading and trailing shroud edges.
[0002] A conventional gas turbine engine includes a core engine having in serial, axial
flow relationship, a high pressure compressor to compress the airflow entering the
core engine, a combustor in which a mixture of fuel and the compressed air is burned
to generate a propulsive gas flow, and a high pressure turbine which is rotated by
the propulsive gas flow and which is connected by a large diameter first shaft to
drive the high pressure compressor. A typical bypass turbofan engine also has a low
pressure turbine aft of the high pressure turbine and which drives a forward fan forward
of the high pressure compressor with a second shaft concentrically disposed within
the first shaft.
[0003] In a typical variable bypass ratio design as disclosed in U.S. Patent No. 4,068,471,
the front fan includes one or more forward rows of fan rotor blades and an aft fan
disposed in serial, axial flow relationship between the forward fan and the high pressure
compressor. The aft fan may have one or more aft rows of fan rotor blades and is connected
to the larger-diameter first drive shaft which is driven by the high pressure turbine.
A variable area bypass injector is located between the forward and aft fans to vary
the amount of air entering a first inlet of a fan bypass duct which varies the fan
bypass ratio of the engine (i.e., the ratio of the air flowing through the fan bypass
duct to the air flowing through the core engine) from which comes the term variable
cycle to describe the engine. The fan bypass duct has a second inlet located aft of
the aft row of fan blades.
[0004] Multiple bypass stream variable cycle gas turbine engines have also been developed
that incorporate flow splitters in the fan and/or compressor sections of the engine,
U.S. Patent Nos. 5,402,638, 5,404,713, and U.S. Patent Application No. 08/624,288
Blisks (one piece bladed integral rotor disc) such as that disclosed in U.S. Patent
No. 5,562,419 have radially inner and outer rows of fan or compressor rotor blade
airfoils that are separated by a rotatable portion of an annular duct wall, often
referred to as a shroud, such that the inner and outer row of rotor blade airfoils
are disposed in separate ducts and flowpaths. The annular duct wall or splitter separates
these two ducts and annular seals are provided between the rotatable portion of the
duct wall and adjacent stationary portions of the duct wall just forward and aft of
the rotatable portion. These blisks may be powered by the high pressure turbine section
and often subjected to very high stress fields due to centrifugal forces during spool
up and high thrust operation. Experience has shown that such a rotor does not do a
good job of transmitting the centrifugal loads from the outer airfoil and shroud into
the inner airfoil. The resulting poor distribution of loading into the inner airfoil
may result in a lack of ability to increase the high pressure compressor rotor speed
to levels that might otherwise be achieved. The shroud must transmit its own centrifugal
loading as well as the outer airfoil centrifugal loading onto the inner airfoil to
be further transmitted into a hub of the blisk. This has a tendency to cause high
stresses at both leading and trailing edges of the inner airfoils. These high stresses
limit the maxim rotational speed at which the rotor could otherwise operate and still
meet good design practices with respect to low and high cycle fatigue.
[0005] Blade designers, therefore, are in need of a rotating flow splitter to significantly
reduce the leading and trailing edge stresses in the inner air foil.
[0006] According to a first aspect of the invention, there is provided a gas turbine engine
blade assembly comprising:
circumferentially disposed and radially extending inner and outer rows of inner and
outer airfoils respectively,
each airfoil having axially spaced apart leading and trailing airfoil edges,
an annular shroud disposed between and connecting said rows of airfoils and having
axially spaced apart leading and trailing shroud edges corresponding to said leading
and trailing airfoil edges respectively, and
an axially extending cavity means in one of said shroud edges for reducing stresses
in corresponding airfoil edges of said inner and outer airfoils wherein said corresponding
airfoil edges are located near said means.
[0007] According to a second aspect of the invention, there is provided a blisk for a gas
turbine engine, said blisk comprising:
an annular rim disposed about a rotor axis,
a blade assembly disposed around said rim,
said blade assembly having circumferentially disposed and radially extending inner
and outer rows of inner and outer airfoils respectively,
each airfoil having axially spaced apart leading and trailing airfoil edges,
an annular shroud disposed between and connecting said rows of airfoils and having
axially spaced apart leading and trailing shroud edges corresponding to said leading
and trailing airfoil edges respectively, and
axially extending cavity means in said leading and trailing shroud edges for reducing
stresses in leading and trailing airfoil edges respectively, and said corresponding
airfoil edges are located near said means.
[0008] The cavities may extend axially under the corresponding airfoil edges. The cavities
may be 360 degree circumferentially extending annular grooves in one embodiment. Another
embodiment has cavities in the form of leading and trailing pluralities of circumferentially
extending annular grooved pockets that axially extend into the leading and trailing
shroud edges, respectively, with each of the pockets being substantially circumferentially
aligned with a corresponding one of the airfoil edges of the outer airfoils. Another
embodiment has a plurality of groups of circumferentially canted holes axially extending
into the shroud edges. Each of the groups is substantially circumferentially aligned
with a corresponding one of the airfoil edges of the outer airfoils and canted in
a direction from the corresponding outer airfoil leading to trailing edges. The blisk
of this invention is particularly useful as part of a rotor powered by a high pressure
turbine of a gas generator in a multiple bypass aircraft gas turbine engine.
[0009] The present invention seeks to provide advantages that include a significant reduction
in the inner airfoil leading and trailing edge stresses caused by the shroud and outer
airfoil centrifugal loads. The ability to run the rotor to higher tip speeds than
with a conventional solid annular shrouds. Under centrifugal loading the addition
of the groove results in additional flexibility of the shroud which causes the leading
and trailing edges of the outer airfoil to be in compression near the airfoil root
which helps provide maximum vibratory stress capability.
[0010] The invention will now be described in greater detail, by way of example, with reference
to the drawings in which:
FIG. 1 is a schematic cross-sectional side view of a bypass turbofan engine incorporating
a blisk having a splitter in accordance with one embodiment of the present invention;
FIG. 2 is an enlarged schematic cross-sectional side view of a more particular embodiment
of the bypass turbofan engine in FIG. 1 around the blisk;
FIG. 3 is an enlarged schematic cross-sectional side view of the blisk in FIG. 2;
FIG. 4 is a perspective view of the blisk in FIG. 3 having an annular groove cavity;
FIG. 5 is a perspective view of a first alternative embodiment of the blisk in FIG.
3 having pocket cavities;
FIG. 6 is a perspective view of a second alternative embodiment of the blisk in FIG.
3 having canted hole cavities; and
FIG. 7 is a perspective view of a third alternative embodiment of the blisk in FIG.
3 having a segmented shroud.
[0011] Referring now to the drawing, there is illustrated in FIGS. 1 and 2 a bypass turbofan
gas turbine engine 10 having a generally longitudinally extending axis or centerline
12 generally extending in a forward direction 14 and an aft direction 16. The bypass
turbofan engine 10 includes a core engine 18 (also called a gas generator) which includes
a core driven fan (CDF) in 19, a high pressure compressor 20, a combustor 22, and
a high pressure turbine (HPT) 23 having a row of high pressure turbine (HPT) blades
24, all arranged in a serial, axial flow relationship. High pressure compressor blades
64 of the high pressure compressor 20 and the CDF 19 are fixedly interconnected in
driving engagement to the high pressure turbine blades 24 by a larger diameter annular
core engine shaft 26 which is disposed coaxially about the centerline 12 of the engine
10 forming a high pressure spool or rotor 29.
[0012] The core engine 18 is effective for generating combustion gases. Pressurized air
from the high pressure compressor 20 is mixed with fuel in the combustor 22 and ignited,
thereby, generating combustion gases. Some work is extracted from these gases by the
high pressure turbine blades 24 which drives the high pressure compressor 20. The
combustion gases are discharged from the core engine 18 into a power turbine or low
pressure turbine (LPT) 27 having a row of low pressure turbine rotor (LPT) blades
28. The low pressure turbine rotor blades 28 are fixedly attached to a smaller diameter
annular low pressure shaft 30 which is disposed coaxially about the centerline 12
of the engine 10 within the core engine shaft 26 forming a low pressure spool. The
low pressure shaft 30 rotates a more longitudinally forward row of generally radially
outwardly extending and circumferentially spaced-apart forward fan rotor blades 32
of a forward fan 33. The core engine shaft 26 also rotates a more longitudinally aft
apart core driven or aft fan rotor blade assembly 36, in accordance with one embodiment
of the present invention, having generally radially outwardly extending blade tips
38. The aft fan rotor blade assembly 36 is disposed longitudinally aft of the more
longitudinally forward row of forward fan rotor blades 32. A row of circumferentially
spaced-apart aft fan stator vanes 34 (attached at either or both radial ends) is disposed
longitudinally between the forward fan 33 and the aft fan rotor assembly 36.
[0013] A fan bypass duct 40 has a first inlet 42 disposed longitudinally between the forward
fan 33 and the aft or core driven fan 19. The first inlet 42 includes a front selector
valve door 44 and a first flow splitter 42A. A second inlet 46 to the fan bypass duct
40 is also disposed longitudinally between the forward fan 33 and the aft or core
driven fan 19, thereby providing two parallel bypass flowpaths into the fan bypass
duct from the forward fan. The fan bypass duct 40 is in fluid communication with a
second inlet 46 by way of a second inlet duct 43 having a second inlet duct outlet
47 to the fan bypass duct 40. The second inlet 46 includes an annular duct wall 45
with a second flow splitter 48. The annular duct wall 45 includes a rotatable portion
commonly referred to as a shroud 108 of the aft CDF 19. Annular seals 45S are provided
between the rotatable shroud 108 of the annular duct wall 45 and adjacent stationary
portions of the wall just forward and aft of it.
[0014] Illustrated in FIG. 2 is a more particular design of the engine 10 and CDF 19 illustrating
the shroud 108 of the annular duct wall 45 that rotates with the high pressure spool.
The second flow splitter 48 may be axially positioned forward of (denoted by the solid
line) or proximate to (denoted by the dashed line) the first flow splitter 42A axial
location depending on the particular aerodynamic considerations for a given engine.
The radially outer airfoils 107 and radially inner airfoils 109 of blade assembly
37 have different airfoils with separate non-continuous profiles and leading and trailing
edges LE and TE, respectively. The outer airfoils 107 may or may not be indexed, one
to one, and generally aligned with the inner airfoils 109. The same construction may
be also be applied to the radially outer vane tip portions 84 and the radially inner
vane hub portions 86 of the fan vanes 34. The vane tip portion 84 may have an extended
length pivotable trailing-edge tip flap 88 that extends aft of the trailing edge 121
of the independently pivotable trailing-edge inner flap 90 radially inner hub portion
86. The annular seals 45S cooperate with lands 138 of adjacent stationary portions
of the wall 45. The engine and its operation are described in more detail in U.S.
Patent Application No. 08/624,288.
[0015] Referring now with more particularity to FIGS. 2 and 3, the shroud 108 is disposed
radially between a radially outer row 116 of outer airfoils 107 and a radially inner
row 111 of inner airfoils 109, respectively of the aft CDF 19. The annular duct wall
45 also includes a non-rotatable portion 106 that is disposed between, preferably
variable angle, radially outer vane tip portions 84 and radially inner vane hub portions
86 of the fan vanes 34. Annular seals 45S are provided to prevent or inhibit leakage
around the shroud 108. The seals 45S include axially spaced apart annular leading
and trailing shroud edges 120 and 122 corresponding to leading and trailing airfoil
edges LE and TE, respectively of the outer airfoils 107. Empty leading and trailing
cavities 128 and 130, respectively extend axially into the shroud 108 from the leading
and trailing shroud edges 120 and 122. This provides a means for reducing stresses
in the leading and trailing airfoil edges LE and TE, respectively of the inner and
outer airfoils 107 and the corresponding airfoil edges are located near the cavities.
The cavities are structural voids that prevent or reduce centrifugal loads from being
transmitted to the leading and trailing airfoil edges LE and TE, respectively of the
inner and outer airfoils 109 and 107, respectively.
[0016] FIGS. 3 and 4 illustrate, in more detail, a more particular embodiment of the present
invention where the blade assembly 36 in the form of a blisk 132 having an annular
hub or rim 140 disposed about a rotor axis which is centerline 12 and the blade assembly
36 is disposed around the rim. The shroud 108 is disposed radially between the radially
outer row 116 of outer airfoils 107 and the radially inner row 111 of inner airfoils
109. The radially outer row 116 of outer airfoils 107 are integrally mounted to the
rim 140. The leading and trailing cavities 128 and 130 are full 360 degree annular
grooves 142 extending axially into the shroud 108 from the leading and trailing shroud
edges 120 and 122. The grooves 142 extend axially under the leading and trailing airfoil
edges LE and TE, respectively of the outer airfoils 107. The blisk of this invention
is particularly useful as part of a rotor powered by the high pressure turbine HPT
of an aircraft gas turbine engine.
[0017] FIG. 5 illustrates another embodiment of the present invention in which the cavities
are in the form of leading and trailing pluralities of circumferentially extending
annular grooved pockets 146 extending axially into the shroud 108 from the leading
and trailing shroud edges 120 and 122. The pockets 146 extend axially under the leading
and trailing airfoil edges LE and TE, respectively of the outer airfoils 107.
[0018] FIG. 6 illustrates another embodiment of the present invention in which the cavities
are in the form of groups 148 of circumferentially canted holes 150 extending axially
into the shroud 108 from the leading and trailing shroud edges 120 and 122. The holes
150 extend axially under the leading and trailing airfoil edges LE and TE, respectively
of the outer airfoils 107. The holes 150 are substantially circumferentially aligned
with a corresponding one of the leading and trailing airfoil edges LE and TE, respectively
of the outer airfoils 107 and canted circumferentially in a direction from the corresponding
outer airfoil leading to trailing edges.
[0019] FIG. 7 illustrates another embodiment of the present invention in which the shroud
108 is segmented such as the one disclosed in U.S. Patent No. 5,562,419
[0020] A friction damper can be designed to fit inside the groove, such as a wire damper,
and used to damp both inner and outer airfoil vibratory modes. The present invention
may also be used with segmented shrouds such as those disclosed in U.S. Patent No.
5,562,419. The annular seals 45S are illustrated as tongue and groove seals where
the tongue is provided by annular knife edges on stationary portions of the annular
duct wall 45 which are cooperatively disposed within the grooves 142. The seals for
the embodiments having the grooved pockets (FIG. 5) and the canted holes (FIG. 6)
use either abutting edges, overlapping edges, or some other kind of sealing means
that are well known in the field such as shiplap seals.
[0021] While the preferred embodiment of our invention has been described fully in order
to explain its principles, it is understood that various modifications or alterations
may be made to the preferred embodiment without departing from the scope of the invention
as set forth in the appended claims.
1. A gas turbine engine blade assembly comprising:
circumferentially disposed and radially extending inner and outer rows of inner and
outer airfoils respectively,
each airfoil having axially spaced apart leading and trailing airfoil edges,
an annular shroud disposed between and connecting said rows of airfoils and having
axially spaced apart leading and trailing shroud edges corresponding to said leading
and trailing airfoil edges respectively, and
an axially extending cavity means in one of said shroud edges for reducing stresses
in corresponding airfoil edges of said inner and outer airfoils wherein said corresponding
airfoil edges are located near said means.
2. An assembly as claimed in claim 1 wherein said cavity means extends axially under
said corresponding airfoil edges.
3. An assembly as claimed in claim 1 or 2, wherein said cavity means is a 360 degree
circumferentially extending annular groove axially extending into said one of said
shroud edges.
4. An assembly as claimed in claim 1 wherein said cavity means is a plurality of circumferentially
extending annular grooved pockets axially extending into said one of said shroud edges
and each of said pockets is substantially circumferentially aligned with a corresponding
one of said airfoil edges of said outer airfoils.
5. An assembly as claimed in claim 1 wherein said cavity means is a plurality of groups
of circumferentially canted holes axially extending into said one of said shroud edges
and each of said groups is substantially circumferentially aligned with a corresponding
one of said airfoil edges of said outer airfoils and canted in a direction from said
corresponding leading to trailing outer airfoil edges.
6. A blisk for a gas turbine engine, said blisk comprising:
an annular rim disposed about a rotor axis,
a blade assembly disposed around said rim,
said blade assembly having circumferentially disposed and radially extending inner
and outer rows of inner and outer airfoils respectively,
each airfoil having axially spaced apart leading and trailing airfoil edges,
an annular shroud disposed between and connecting said rows of airfoils and having
axially spaced apart leading and trailing shroud edges corresponding to said leading
and trailing airfoil edges respectively, and
axially extending cavity means in said leading and trailing shroud edges for reducing
stresses in leading and trailing airfoil edges respectively, and said corresponding
airfoil edges-are located near said means.
7. A blisk as claimed in claim 6 wherein said cavity means extends axially under said
corresponding airfoil edges.
8. A blisk as claimed in claim 6 or 7 wherein said cavity means are 360 degree circumferentially
extending annular grooves, one of each of said grooves extending axially into one
of each of said shroud edges.
9. A blisk as claimed in claim 6 or 7 wherein said cavity means is leading and trailing
pluralities of circumferentially extending annular grooved pockets axially extending
into said leading and trailing shroud edges respectively and each of said pockets
is substantially circumferentially aligned with a corresponding one of said airfoil
edges of said outer airfoils.
10. A blisk as claimed in claim 6 or 7 wherein each of said cavity means is a plurality
of groups of circumferentially canted holes axially extending into said shroud edges
and each of said groups is substantially circumferentially aligned with a corresponding
one of said airfoil edges of said outer airfoils and canted in a direction from said
corresponding leading to trailing outer airfoil edges.